It has become increasingly important to eliminate the features associated with a ballistic trajectory ordinarily followed by rockets and other jet-propelled projectiles, by forming the projectiles as spherical spin-stabilized missiles. The spherical missile spins about an axis upwardly inclined relative to the intended straight line path of flight and aligned with the missile propulsion thrust axis. The missile is released following ignition or activation of the propulsion system within the missile. The propulsion is effected by the reaction of the exhaust jet of, for example, a rocket motor housed within the spherical missile shell.
Often such spherical spin-stabilized missiles are provided in conjunction with attachments secured to the front end of an assault weapon such as a rifle.
Such spin-stabilized spherical self-propelled missiles experience difficulties in remaining stabilized during attainment of desired rotational speed and in coordinating the spinning and release of the missile. Release of the missile prior to attainment of adequate rotational speed can result in unstable flight. Delay of release after attainment of adequate rotational speed can result in a loss of propulsive range.
Consequently, attempts have been made to provide means for temporarily restraining and automatically releasing a spin-stabilized self-propelled spherical missile during spinup. For instance, in U.S. Pat. No. 3,245,350 to J. A. Kelly, dated Apr. 12, 1966, a mechanical release is provided between a rifle barrel and a spin-stabilized spherical missile in order to selectively release the missile. However, precise automatic release is not afforded. More specifically, U.S. Pat. No. 3,554,078 to Joseph S. Horvath, dated Jan. 12, 1971, provides a fusible link for temporarily restraining and automatically releasing a spherical spin-stabilized missile during spinup. Release of the spherical rocket missile from its rotary supporting means is effected by causing hot missile rocket exhaust gas to weaken by heating or to heat and soften or melt a separate fusible link member which, prior to weakening by softening or melting, secures the missile to the rotary support means. In this patent, the separate fusible link member is of the nature of a brazing alloy serving as one part of a nozzle assembly to secure the rocket to the rotary support means. The fusible link member is brazed between two separate fore and aft nozzle portions which are permanently secured to the missile and to the support means, respectively, as by threaded engagements.
An improvement on the aforementioned prior art is disclosed in U.S. Pat. No. 4,395,836 to Baker et al, dated Aug. 2, 1983 and assigned to the assignee of this invention, wherein a new and improved nozzle assembly is disclosed. The nozzle assembly includes a unitary nozzle member having fusible joint means formed integrally therewith, between the missile and the rotary support means, thereby eliminating the assembly and brazing operations of prior devices as shown in the Horvath patent, and thereby considerably reducing manufacturing costs and improving accuracy. However, in this patent the fore and aft sections of the unitary nozzle, forwardly and rearwardly of the fusible joint means, are permanently fixed to the missile and to the support means, respectively, as by threaded engagements.
Further improvements are shown in U.S. Pat. No. 4,403,435 to Baker et al, dated Sept. 13, 1983 and assigned to the assignee of this invention, wherein a further new and improved nozzle assembly includes projectile support means having open-ended receptacle means out of which fore and aft sections of the nozzle can move on fusing and separation of the fusible joint means. This patent also shows an improved register section for the missile or nozzle which is generally conical in configuration to improve alignment of the missile with the spin axis during initial separation of the fusible joint means.
The present invention represents somewhat of a radical departure from the prior art in that a mass is caused to be urged or propelled rearwardly by the gases of the missile or separate or combined other force generating mechanism to strike an abutment means on the turbine or rotary means for the missile to cause the rotary means in its receptacle, to move rapidly away from the missile after separation of the fusible joint means. The present invention thus allows positive missile retention by the launch system rotary means during coupling fusing and therefore eliminates pointing error tip off forces initiated by the coupling fusing in any of the prior art.